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National Advisory Committee for Aeronautics 



N0.93 



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esearc 



h Abstracts 



NOVEMBER 30 



CURRENT NACA REPORTS 

NACA RM E55B11 

FULL-SCALE PERFORMANCE STUDY OF A PRO- 
TOTYPE CRASH-FIRE PROTECTION SYSTEM FOR 
RECIPROCATING-ENGINE-POWERED AIRPLANES. 
Dugald O. Black and Jacob C. Moser. November 
1955. 36p. diagrs., photos. (NACA RM E55B11) 

An airplane was experimentally crashed to study the 
performance of a prototype crash-fire inerting sys- 
tem for reciprocating-engine-powered airplanes. 
The results of previous experimental crashes indi- 
cate that the crash conditions imposed almost always 
result in fire. The inerting system was therefore 
exposed to conditions that would adequately test its 
ability to inert and de-energize the various ignition 
sources known to cause crash fires. The fact that 
fire did not occur during this crash indicated that the 
crash-fire inerting system functioned satisfactorily 
as a complete unit. The prototype inerting system 
functioned with a rapidity equal to or greater than 
that of the experimental system used in the NACA 
crash-fire studies. 



NACA RM E55H11 

A SURVEY OF UNCLASSIFIED AXIAL -FLOW- 
COMPRESSOR LITERATURE. Howard Z. Herzig 
and Arthur G. Hansen. November 1955. i, 88p. 
(NACA RM E55H11) 

A survey of unclassified axial-flow-compressor 
literature is presented in the form of brief reviews 
of the methods, results, and conclusions of selected 
reports. The reports are organized into several 
main categories with subdivisions, and frequent ref- 
erences are made within the individual reviews to 
pertinent material elsewhere in the survey. 



NACA RM E55127a 

AVERAGE BOND ENERGIES BETWEEN BORON 
AND ELEMENTS OF THE FOURTH, FIFTH, SIXTH, 
AND SEVENTH GROUPS OF THE PERIODIC TABLE. 
Aubrey P. AltshuUer. November 1955. 7p. tab. 
(NACA RM E55I27a) 

The average bond energies D (B-Z) for boron- 
containing molecules have been calculated by the 
Pauling geometric-mean equation. These calculated 
bond energies are compared with the average bond 
energies Dg (B-Z) obtained from experimental 
data. Thehigher values of Dgxp(B-Z) in compar- 
ison with lUijj(B-Z) when Z is an element in the 
fifth, sixth, or seventh periodic group may be attrib- 
uted to resonance stabilization or double-bond char- 
acter. 




NACA TM 1384 

METALLOGRAPHY OF ALUMINUM ANIVITi 
ALLOYS. USE OF ELECTROLYTIC POLiSH 
(Metallographie de I'aluminium et de ses alHages^ 
Emploi du polissage electrolytique). P. A. Jafrquet. 
November 1955. ii, BOp. diagrs., photos., tabs. " 
(NACA TM 1384. Trans, from Office National 
d'Etudes et de Recherches Aeronautiques, Pub. 51, 
1952) 

Recent methods are described for electropolishing 
aluminum and aluminum alloys. Numerous refer- 
ences are included of electrolytic micrographic 
investigations carried out during the period 1948 to 
1952. A detailed description-af a commercial elg§-' 
trolytic polishing unit, suitable for micrographic Ex- 
amination of aluminum and its alloys, is included. 



NACA TN 3413 



INVESTIGATION OF THE tJSfiOF A RUBBER ANA- 
LOG IN THE STUDY OF STRESS DISTRIBUTION IN 
RIVETED AND CEMENTED JOINTS. Louis R. 
Demarkles, Massachusetts Institute of Technology. 
November 1955. 97p. diagrs., tabs. 
(NACA TN 3413) 

Results are presented of an investigation made to 
study the stress distribution within cemented and 
riveted joints by use of an analogous joint con- 
structed of a highly flexible material. Displacement 
measurements obtained from foam-rubber analogs, 
and rational though not rigorously sound formulas 
for shear stress distribution in joints, are given. 



NACA TN 3469 

SUMMARY OF RESULTS OBTAINED BY 
TRANSONIC-BUMP METHOD ON EFFECTS OF 
PLAN FORM AND THICKNESS ON LIFT AND DRAG 
CHARACTERISTICS OF WINGS AT TRANSONIC 
SPEEDS. Edward C. Polhamus. November 1955. 
33p. diagrs., tab. (NACA TN 3469. Supersedes 
RM L51H30) 

This paper presents a summary of the effects of 
plan form and thickness on the lift and drag charac- 
teristics of wings at transonic speeds and compari- 
sons with subsonic, transonic, and supersonic 
theories. The data considered in this summary were 
obtained during a transonic research program con- 
ducted in the Langley high-speed 7- by 10-foot tunnel 
by the transonic-bump method. The Reynolds num- 
bers of the tests were generally less than 1 x 10^. 



I 

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U.S. DEPOWT OBtV f 

II 11 ■■■ HIT" 



•AVAILABLE ON LOAN ONLY. 

ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 15H H ST., NW., WASHINGTON 25, D C, CITING CODE NUMBER ABOVE EACH TITLE; 

THE REPORT TITLE AND AUTHOR. 



NACA TN 3492 

DETERMINATION OF INFLOW DISTRIBUTIONS 
FROM EXPERIMENTAL AERODYNAMIC LOADING 
AND BLADE-MOTION DATA ON A MODEL HELI- 
COPTER ROTOR IN HOVERING AND FORWARD 
FLIGHT. Gaetano Falabella, Jr., and John R. 
Meyer, Jr.. Massachusetts Institute of Technology. 
November 1955. 184p. diagrs., photos., tab. 
(NACA TN 3492) 

Inflow distributions, azimuth and spanwise, were 
determined analytically from measured pressure 
distributions and blade-motion data on a model heli- 
copter rotor blade under hovering and simulated 
forward-flight conditions. Pressures and corre- 
sponding blade flapping were recorded for various 
rotor conditions at tip-speed ratios of 0.10 to 1.00. 
Supplementary information concerning reverse-flow 
effects on offset -blade motion, measured forces and 
moments on a typical offset model rotor, and addi- 
tional recorded pressure data are also included. 



NACA TN 3524 

THE EFFECT OF REYNOLDS NUMBER ON THE 
STALLING CHARACTERISTICS AND PRESSURE 
DISTRIBUTIONS OF FOUR MODERATELY THIN 
AIRFOIL SECTIONS. George B. McCuUough. 
November 1955. 24p. diagrs., tabs. 
(NACA TN 3524) 

Low-speed measurements of the lift, drag, pitching 
moment, and pressure distribution of the NACA 
0008, 0007.5, 0007, and 0006 airfoil sections are 
presented for Reynolds numbers from 1.5 to 6 mil- 
lion. It is shown that the flow over these sections 
underwent a change at some value of the lift coeffi- 
cient which depended on the airfoil thickness ratio 
and Reynolds number. The effect of the flow change 
on maximum lift was small. 



NACA TN 3525 

VORTEX INTERFERENCE ON SLENDER AIR- 
PLANES. Alvin H. Sacks. November 1955. 
diagr. (NACA TN 3525) 



19p. 



Formulas are developed for the forces and moments 
due to vortex interference on a slender wing-body- 
tail combination of general cross section performing 
quasi-stationary maneuvers. It is found that in 
steady straight flight the interference lift depends 
only on the impulse of each shed vortex and its 
image vortex in a transformed circle plane, this 
quantity to be determined at the wing trailing edge 
and at the base of the configuration. 



NACA 
RESEARCH 



ABSTRACTS NO. 93 



NACA TN 3526 

FLIGHT CALIBRATION OF FOUR AIRSPEED SYS- 
TEMS ON A SWEPT-WING AIRPLANE AT MACH 
NUMBERS UP TO 1.04 BY THE NACA RADAR- 
PHOTOTHEODOLITE METHOD. Jim Rogers 
Thompson, Richards. Bray, and George E . Cooper. 
November 1955. 41p. diagrs., photos., tab. 
(NACA TN 3526. Supersedes RM A50H24) 

The characteristics of four different airspeed sys- 
tems installed in a swept-wing airplane have been 
investigated in flight up to 1.04 Mach number by the 
NACA radar-phototheodolite method of airspeed 
calibration. The variations of static-pressure de- 
fect per unit indicated impact pressure with Mach 
number and a limited amount of information on the 
effect of airplane normal-force coefficient are pre- 
sented for each system. The results are compared 
with available theory and wind-tunnel tests of the 
isolated heads. 



NACA TN 3547 

AERODYNAMIC CHARACTERISTICS OF A SMALL- 
SCALE SHROUDED PROPELLER AT ANGLES OF 
ATTACK FROM 0° TO 90°. Lysle P. Parlett. 
November 1955. 12p. diagrs. (NACA TN 3547) 

Tests have been performed to determine the effects 
of airspeed and angle of attack on the lift, drag, and 
pitching moment of a shrouded-propeller model, 
having a shroud length of about two-thirds of the 
propeller diameter, over an angle-of-attack range 
from 0° to 90°. Tests were made of the complete 
model with the propeller operating and also of the 
shroud alone with the propeller removed. The effect 
of inlet-lip cross-sectional radius on the static- 
thrust characteristics was also studied. 



NACA TN 3548 

FLIGHT INVESTIGATION AT MACH NUMBERS 
FROM 0.6 TO 1.7 TO DETERMINE DRAG AND 
BASE PRESSURES ON A BLUNT-TRAILING-EDGE 
AIRFOIL AND DRAG OF DIAMOND AND CIRCULAR- 
ARC AIRFOILS AT ZERO LIFT. John D. Morrow 
and Ellis Katz. November 1955. 19p. diagrs., 
photos. (NACA TN 3548. Supersedes RM L50E19a) 

Results of an exploratory free-flight investigation at 
zero lift of several rocket-powered drag-research 
models having rectangular 6-percent -thick wings are 
presented for a Mach number range of 0.6 to 1.7. 
Wings having diamond, circular-arc, and blunt- 
trailing-edge airfoil sections were tested. Pres- 
sures over the base of the blunt -trailing-edge airfoil 
were measured. The drags of all the models were 
measured and are compared with theory in this 
paper. 



NACA 
RESEARCH 



ABSTRACTS NO. 93 



NACA TN 3569 

COMPRESSIBLE LAMINAR BOUNDARY LAYER 
AND HEAT TRANSFER FOR UNSTEADY MOTIONS 
OF A FLAT PLATE. Simon Ostrach. November 
1955. 26p. diagrs., tab. (NACA TN 3569) 

The laminar compressible boundary layer and heat 
transfer over an isothermal semi-infinite flat plate 
moving with a time-dependent velocity has been ana- 
lyzed. First-order deviations from the quasi- 
steady velocity and temperature profiles and 
boundary-layer characteristics have been computed. 
A plate oscillating about a steady velocity is consid- 
ered as an example. 



NACA TN 3574 

ACOUSTIC ANALYSIS OF RAM-JET BUZZ. Harold 
Mirels. November 1955. 33p diagrs. 
(NACA TN 3574) 

A one-dimensional analysis of ram-jet buzz is pre- 
sented. It is assumed that the buzz has a linear 
instability origin and that the combustion chamber is 
of constant area. The configuration is shown to be 
unstable when the real part of the acoustic impedance 
of the inlet is greater than a term of the order of the 
combustion-chamber Mach number. Computations 
indicate that burning with a fixed planar flame front 
and constant heat release per unit mass increases 
the stable operating range. 



NACA TN 3583 

CHARTS OF BOUNDARY-LAYER MASS FLOW AND 
MOMENTUM FOR INLET PERFORMANCE ANALY- 
SIS MACH NUMBER RANGE. 0.2 TO 5.0. Paul C. 
Simon and Kenneth L. Kowalski. November 1955. 
32p. diagrs., tab. (NACA TN 3583) 

Significant flow parameters for various fractions of 
a turbulent boundary layer are presented in chart 
form for a number of power-law velocity profiles 
and a range of Mach numbers up to 5.0. Estimates 
of auxiliary inlet mass flow or momentum may easily 
be made. Application of the charts to inlets of 
arbitrary shape and to the determination of the pres- 
sure recovery of rectangular normal-shock inlets 
immersed in boundary layer is described. 



NACA TN 3586 

IMPINGEMENT OF WATER DROPLETS ON NACA 
65A004 AIRFOIL AT 0° ANGLE OF ATTACK. 
Rinaldo J. Brun and Dorothea E. Vogt. November 
1955. 28p. diagrs. (NACA TN 3586) 

The trajectories of droplets in the air [lowing past 
an NACA 65A004 airfoil at an angle of attack of 0° 
were determined. The amount of water in droplet 
form impinging on the airfoil, the area of droplet 
impingement, and the rate of droplet impingement 
per unit area on the airfoil surface were calculated 
from the trajectories and presented to cover a large 
range of flight and atmospheric conditions. These 
impingement characteristics are compared briefly 
with those previously reported for the same airfoil 
at angles of attack of 4° and 8°. 



BRITISH REPORTS 



N-40040* 

National Gas Turbine Establishment (Gt. Brit.) 
AN EXPERIMENTAL INTRODUCTION TO THE JET 
FLAP; N. A. Dimmock. July 1955. 68p. diagrs., 
photos., tabs. (NGTE R. 175) 

Results are given of two airfoils, each having 12.5- 
percent-thick elliptical cross section with a narrow 
full-span jet slot at the trailing edge, the jet deflec- 
tions being, respectively, 90° and 31.4°. The val- 
ues of the force and moment coefficients and deriv- 
atives agree with those suggested by theory in a pre- 
vious report. Support is given to the thrust hypoth- 
esis in that the measured thrust was greater, under 
appropriate conditions, than the reaction component 
from the deflected jet. The losses in the system 
are considered and some are investigated, those due 
to Reynolds number and jet entrainment effects being 
included. Influence of ground on lift and center of 
lift was measured and found not to be prohibitive. 



N-40048 

Royal Aircraft Establishment (Gt. Brit. ) 
TESTS OF HUMIDITY EFFECTS ON FLOW IN A 
WIND TUNNEL AT MACH NUMBERS BETWEEN 
2.48 AND 4. R. J. Monaghan. January 1955. 33p. 
diagrs. (RAE Tech. Note Aero 2358) 

Static and pilot pressure distributions were meas- 
ured in the working section of a 5-in. by 5-in. 
supersonic wind tunnel at nominal Mach numbers of 
2.48, 3.25, and 4, over ranges of absolute humidity 
at the inlet from 5 x 10"^ to 3 x 10'^. Previous 
work indicates that a condensation shock would occur 
in the nozzle. For a stagnation pressure of 1 at- 
mosphere and stagnation temperatures giving zero 
heat-transfer conditions at the walls, no humidity 
effects were discernible if the absolute humidity was 
less than 2 x 10"^ at M = 2.48, 3 x 10"* at M = 3.12, 
and about 5 x 10"* at M = 3.8. 



N-40068 

Aeronautical Research Council (Gt. Brit.) 
TESTS ON A SWEPT-BACK WING AND BODY WITH 
ENDPLATES AND WING TIP TANKS IN THE COM- 
PRESSED AIR TUNNEL. C. Salter and R. Jones. 
APPENDIX - COMPARISON BETWEEN THE MEAS- 
URED LIFT AND DRAG AND CALCULATED VAL- 
UES FOR THE WING WITH TIP TANKS. J. Weber. 
1954. 26p. diagrs.. tabs. (ARC CP 196) 

Results are given of experiments to determine the 
effect on lift, drag and pitching moment, of wing tip 
tanks and of two sizes of end plates on a tapered 
swept wing model. The tests were undertaken pri- 
marily to extend the range of Reynolds number for 
checks on previous theoretical work. As regards 
lift and pitching moment, the effects are found to be 
fairly well defined. The drag characteristics are 
less consistent, but it seems that end plates have 
the effect of reducing the drag of the model over 
quite a large range of C^- This does not apply in 
the case of the wing lip tanks. 



N-40069' 

Aeronautical Research Council (Gt. Brit. ) 
WIDE RANGE AMPLIFIER FOR TURBULENCE 
MEASUREMENTS WITH ADJUSTABLE UPPER 
FREQUENCY LIMIT. H. Schuh and D. Walker. 
1955. 42p. diagrs. (ARC CP 198) 

Requirements are discussed for an amplifier suit- 
able for subsonic and supersonic turbulence work 
with hot wires. An amplifier is described which has 
a frequency range from 1.4 c/s to 50 kc/s, dealing 
with a range of thermal time lag from 0.1 m. s. to 
5 m. s. An iron dust-cored inductance is used to 
give the required compensation for thermal lag, the 
circuit being a modification of Dryden's circuit. 
The upper frequency cutoff is adjustable in six steps 
from 1.5 kc/s to 50 kc/s. The output can be applied 
to a thermocouple meter and to an oscilloscope. 



N-40080'' 

Aeronautical Research Council (Gt. Brit. ) 
FLIGHT TESTS AT TRANSONIC SPEEDS ON FREE- 
LY FALLING MODELS. PARTS I TO V. Edited by 
C. Kell. PART I - HISTORICAL. C. Kell. PART 
n - EQUIPMENT AND TECHNIQUE. C. Kell and 
J. Swan. PART III -DRAG EXPERIMENTS. 
C. Kell and F. Smith. PART IV - FLUTTER EX- 
PERIMENTS. W. G. Molyneux and E. W. Chappie. 
PART V - NOTES ON THE ACCURACY OF THE 
FREELY FALLING MODEL EXPERIMENT. 
T. F. C. Lawrence. 1955. 32p. diagrs., photos., 
tab. (ARC R fe M 2902. Supersedes RAE Tech. 
Memo. Aero 308) 

Basic bodies carrying the airfoils to be tested were 
released from an aircraft flying at height, and accel- 
erated under the influence of gravity through the 
transonic speed range. Radar recorded the flight 
path and telemetering equipment carried within the 
body transmitted information to a ground station 
during the free fall. This work started in 1943 and 
was brought to a close in 1949. 



N-40082* 

Aeronautical Research Council (Gt. Brit.) 
OBSERVATIONS OF THE FLOW ROUND A TWO- 
DIMENSIONAL AEROFOIL OSCILLATING IN A 
HIGH-SPEED AIR STREAM. A. Chinneck, D. W. 
Holder, and C. J. Berry. 1955. 18p. diagrs., 
photos., tab, (ARC R i: M 2931. Supersedes ARC 
15. 141; FM 1779 & 0.1006) 

Photographs have been taken of the flow around a 
10-percent-thick RAE 104 airfoil performing pitch- 
ing oscillations at low values of the frequency pa- 
rameter in subsonic and supersonic airstreams. 
Apart from a difference of phase, the general flow 
pattern appeared to be similar to those observed in 
steady motion, the pattern for a particular instanta- 
neous incidence of the oscillation resembling that 
for steady motion at a different incidence. It is 
suggested that, for the range of frequency parameter 
covered, the observed phase lag of the flow pattern 
corresponds to the lag in the circulation. 



NACA 

RESEARCH ABSTRACTS NO. 93 



N-40084 

Aeronautical Research Council (Gt. Brit.) 
THE AERODYNAMIC EFFECTS OF ASPECT RATIO 
ON FLUTTER OF UNSWEPT WINGS. W. G. 
Molyneux and E. W. Chappie. 1955. 12p. diagrs., 
tab. (ARC R S; M 2942; 15,609. Supersedes RAe' 
Structures 135) 

A method is described for the direct measurement o£ 
the aerodynamic effects of aspect ratio on wing flut- 
ter. The method requires the use of stiff (virtually 
rigid) wings fle.xibly mounted at the root. Details 
are given of tests on untapered, unswept wings with 
freedoms in modes of linear flexure and uniform 
pitch. A comparison is made between measured 
values of the flutter characteristics and the values 
calculated using an aerodynamic theory for oscil- 
lating wings of finite aspect ratio, and reasonable 
agreement for flutter speeds and frequencies is 
obtained. 



N-40186* 

Royal Aircraft Establishment (Gt. Brit.) 
MEASUREMENTS OF PITCHING MOMENT DERIVA- 
TIVES FOR A SERIES OF RECTANGULAR WINGS 
AT LOW WIND SPEEDS. P. R. Guyett and D. E. G. 
Poulter. June 1955. 52p. diagrs., tabs. (RAE 
Structures 185) 

The direct aerodynamic moments for pitching oscil- 
lations have been measured on a series of rectangu- 
lar wings having aspect ratios between 2 and 8 for 
axis positions at the wing leading edges and trailing 
edges. Two of the wings were also tested with 
single end plates which were aerodyna.mically effec- 
tive in doubling the wing geometric aspect ratio. 
The measurements were made at low speeds in an 
open jet wind tunnel ajid covered the range of fre- 
quency parameter (based on wing chord) 0.13 to 0.39. 
The results are in general agreement with theoreti- 
cal results due to Lawrence and Gerber. Similar 
tests were also made on a wing fitted with two end 
plates in an attempt to obtain results for two- 
dimensional flow. The results do not agree with 
other experimental results and two-dimensional 
theoretical values and indicate that wind-tunnel in- 
terference is important for this test configuration. 



N-40187 

Royal Aircraft Establishment (Gt. Brit.) 
THE MECHANICAL PROPERTIES AND STRUCTURE 
OF CONTINUOUSLY CAST A.C.9 ALUMINIUM 
ALLOY TUBES. P. C. Bradley and D. Bunting. 
June 1955. lOp. diagr., photos., labs. (RAE Tech. 
Note RPD 122) 

The mechanical properties of a large diameter tube 
continuously cast in aluminum alloy A.C. 9 have been 
investigated at room and at elevated temperature, in 
connection with the use of such tubes as fuel or 
oxidant tanks for rocket motors. The variation in 
structure of the alloy across a section is described 
and discussed with reference to the possible detri- 
mental effect of primary silicon crystals on the 
mechanical properties. The operating temperature 
should not be allowed to exceed about 180° C if tanks 
of this material are to be used more than once. 



NACA 

RESEARCH ABSTRACTS NO. 93 



N-40188'' 



Royal Aircraft Establishment (Gt. Brit.) 
THE EFFECTS OF TAPER ON THE SUPERVELOC- 
ITIES ON THREE-DIMENSIONAL WINGS AT ZERO 
INCIDENCE. K. W. Newby. June 1955. 120p. 
diagrs., tab. (RAE Aero 2544) 

Relationships have been derived for expressing the 
velocities on three-dimensional tapered wings at 
zero incidence in terms of the velocities on un- 
tapered infinite swept wings. The theoretical inves- 
tigation of the effects of taper is confined to simple 
wings having airfoil sections formed by cubic or 
parabolic arcs; some exj^erimental evidence is given 
to show that the results of this investigation can 
probably be applied quantitatively to wings having 
conventional airfoil sections. The results given in 
this report show that plan form and thickness taper 
have a marked effect on the velocities near the cen- 
ter of a wing, but that these effects decrease with in- 
crease of sweepback. A calculation method is out- 
lined in section 4.26 of the text for applying the re- 
sults obtained for wings having parabolic arc airfoil 
sections, to wings having arbitrary section shapes. 



N-40189* 

Royal Aircraft Estaclishment (Gt. Brit.) 
RECORDING AND PROCESSING FLIGHT TEST 
DATA BY DIGITAL METHODS. E. J. Petherick. 
(Prepared for AGARD Flight Test Panel). April 
1955. 12p. photos. (RAE Tech. Note MS 20) 

This note reviews current developments in the re- 
cording and processing of flight test data by digital 
methods. It first describes four assessors which 
facilitate conversion of strip chart and kinetheodolite 
records to typewritten or punched card form. It 
then details some coded scales which can be read 
automatically, and finally it describes the incorpora- 
tion and use of such scales in a digital recording 
system intended for airborne use. 



N-40190* 

Royal Aircraft Establishment (Gt. Brit.) 
WIND TUNNEL TESTS ON A 6 FT DIAMETER HEL- 
ICOPTER ROTOR. T. B. Owen, R. A. Fail, and 
R. C. W. Eyre. May 1955. 33p. diagrs., tabs. 
(RAE Tech. Note Aero 2378) 

Thrust, torque, and flapping angle have been meas- 
ured on a 6-foot diameter rotor over a range of 
blade angle, shaft inclination, and tip speed ratio for 
comparison with the 12-foot diameter rotor previ- 
ously tested in the 24-foot tunnel. In addition to 
tests in the 24-foot tunnel, the 6-foot diameter rotor 
was also tested in the No. 2 11-1/2 foot tunnel to 
investigate tunnel constraint. Brief investigations 
were made of support interference and blade twist- 
ing. There are small discrepancies both as regards 
tunnel corrections and as regards the comparison of 
the 6-foot and 12-foot diameter rotors in the un- 
stalled operating range. Possible reasons are dis- 
cussed but small unexplained discrepancies remain. 
Blade stalling has larger effects on the 6- foot diam- 
eter rotor but owing to the progressive nature of the 
phenomenon it is not possible to define any precise 
limits to the ranges of validity of the results on the 
two rotors. 



N-40191* 

Aeroplane and Armament Experimental Establish- 
ment (Gt. Brit.) AN EXPERIMENTAL INVESTIGA- 
TION INTO THE PERFORMANCE OF A HELICOP- 
TER FOLLOWING SUDDEN REDUCTION IN POWER. 
G. W. Langdon. August 4, 1955. 12p. diagrs. 
(AAEE/Res/289). 

The performance of a single rotor helicopter fitted 
with a throttle override has been measured under 
conditions simulating the failure of one engine of a 
multiengined helicopter during take-off. Records of 
the motion are included and the experimental results 
are compared with theoretical predictions. 



N-40192 

Royal Aircraft Establishment (Gt. Brit.) 
INVESTIGATION OF THE FATIGUE OF EXTRUDED 
TUBULAR BOOMS. W. A. P. Fisher and 
H. Yeomans. June 1955. 15p. diagrs., photos., 
tabs. (RAE Tech. Note Structures 162). 

The presence of the unmachined extruded surface of 
the bore in the tubes, as used for the "Viking" and 
"Valetta" spar booms, has a marked adverse effect 
on the basic fatigue strength of the tube. The fail- 
ures of the necked specimens show that the flaws at 
the inner surface are a source of fatigue. Scatter in 
the endurance of tubular spar booms is probably 
largely due to the chances of such flaws occurring at 
the side of a transverse hole. The specimens were 
made from aluminum alloy extruded tube DTD 364. 



N-40193* 

Marine Aircraft Experimental Establishment. (Gt. 
Brit.) INVESTIGATION OF HIGH LENGTH BEAM 
RATIO SEAPLANE HULLS WITH HIGH BEAM 
LOADINGS - HYDRODYNAMIC STABILITY. 
PART 17 - THE STABILITY AND SPRAY CHARAC- 
TERISTICS OF MODEL M. J. K. Fris\vell, D. M. 
Ridland, and A. G. Kurn. April 1955. 24p. diagrs., 
photos., tabs. (MAEE F Res 253) 

In this report results are presented of limited tests 
on the hydrodynamic characteristics of model M of 
the series, tliese tests being designed solely to pro- 
vide information on the interactions of the different 
relevant parameters. The model has a length-to- 
beam ratio of 13 (the forebody being 6 beams in 
length and the afterbody 7 beams), no forebody warp, 
an afterbody to forebody keel angle of 8'^, and a 
straight transverse step with a step depth of 0.15 
beams. The tests comprised the determination of 
longitudinal stability limits without slipstream at 

C. =2.75 and an investigation of spray at this load- 

"o 
ing. A short discussion of the results is also 

included. 



N-40197* 

Royal Aircraft Establishment (Gt. Brit.) 
CORRELATED FATIGUE DATA FOR AIRCRAFT 
STRUCTURAL JOINTS. R. B. Heywood. June 1955. 
16p. diagrs., tab. (RAE Structures 184) 

Results of fatigue tests carried out at RAE on typi- 
cal aircraft wing structural joints are correlated to 
give an indication of general fatigue behavior. The 
results are plotted in the form of S - Log N curves, 
and these indicate that the mode of behavior cannot 
be attributed to any single factor, such as the type of 
aluminum alloy, the ultimate tensile strength, or the 
mean stress of the fatigue cycle. The detailed meth- 
od of design undoubtedly has a predominant influence 
on behavior, but this quality is not revealed by a 
broad classification according to the proportion of 
load transmitted at holes. 



N-40198* 

Royal Aircraft Establishment (Gt. Brit.) 
A TECHNIQUE FOR THE MEASUREMENT OF 
PRESSURE DISTRIBUTION ON OSCILLATING 
AEROFOILS, WITH RESULTS FOR A RECTANGU- 
LAR WING OF ASPECT RATIO 3.3. W. G. Molyneux 
and F. Ruddlesden. June 1955. 24p. diagrs., tabs. 
(RAE Tech. Note Structures 164) 

Details are given of a strain-gage pressure trans- 
ducer that has been developed for measurements of 
pressure distribution on oscillating airfoils in low 
speed wind tunnels. The transducer characteristics 
are shown to be well suited to oscillatory measure- 
ments, and in particular the transducer output can be 
measured directly on a sensitive galvanometer with- 
out the need for preamplification. As an illustration 
of the use of the transducer, pressure measurements 
have been made in the RAE 5 -foot diameter open jet 
wind tunnel on a rectangular wing of aspect ratio 3.3 
oscillating in modes of pitch and roll. Values for 
the aerodynamic derivatives have been obtained from 
the integrated pressure distributions, and are com- 
pared with those derived from overall force meas- 
urements and with theoretical values. The measured 
values are in close agreement but there are some 
discrepancies with theory that are thought to be due 
to a wind-tunnel interference effect. 



N-40199* 

Royal Aircraft Establishment (Gt. Brit.) 
THE INFLUENCE OF PRE-LOADING ON THE FA- 
TIGUE LIFE OF AIRCRAFT COMPONENTS AND 
STRUCTURES. R. B. Heywood. June 1955. 27p. 
diagrs., photos., tabs. (RAE Structures 182) 

Tests on aircraft components and structures are 
described which show that preloading can have a 



NACA 

RESEARCH ABSTRACTS NO. 93 



marked influence on fatigue behavior. Tensile pre- 
loading may increase the life - in one instance a 
hundredfold improvement was obtained - and com- 
pressive preloading may reduce the life. The effect 
is attributed to residual stresses and to load re- 
distributions induced by preloading. 



N-40205* 

Ministry of Supply (Gt. Brit.) 

AN ABSORPTIOMETRIC METHOD FOR THE DE- 
TERMINATION OF CHROMIUM IN 
METHACRYLATO-CHROMIUM TREATED GLASS 
FABRIC. E. J. McLauchlan. August 1955. 4p. 
diagr. (MOS AID Chem. 6) 

An absorptiometric method for the determination of 
chromium in methacrylato-chromium treated glass 
fabric is described. A wet-oxidation attack is uti- 
lized to remove the chromium from the glass fibers 
and to oxidize the tervalent chromium to sexavalent 
chromium for subsequent absorptiometric deter- 
mination as the diphenylcarbazide complex. 



N-40207* 

Royal Aircraft Establishment (Gt. Brit.) 
OPTIMUM DESIGNS FOR REINFORCED CIRCULAR 
HOLES. E. H. Mansfield. June 1955. 26p. diagrs. 
(RAE Structures 183) 

The design of reinforced circidar holes in an Infinite 
sheet is considered theoretically. The stress sys- 
tem in the main body of the sheet is assumed to be 
one in which the principal stresses are in the ratio 
1:-1 (that is, shear), 1:0 (that is, tension), 1:1 or 
1:1 '2. The reinforcement may vary round the hole 
and families of such reinforcements with constant 
total weight are considered; the peak stresses in the 
sheet are evaluated so that optimum weight -strength 
designs are determined. 



N-40209 

Royal Aircraft Establishment (Gt. Brit.) 
STATIC ELECTRICITY EN AIRCRAFT FUEL TANKS. 
F. L. Holmes and D. T. Sharwood. May 1955. lOp. 
diagrs., tab. (RAE Tech. Note EL. 82; Tech. Note 
Mech. Eng. 201) 

Tests were made to measure the static field strength 
generated in aircraft fuel tanks by foaming of the 
fuel. A type of apparatus for measuring static field 
strength is described with details of its application 
to these tests. The results show that the field 
strength is less than 1 volt per centimeter which is 
many times less than that required for flashover and 
the ignition of a fuel vapor/ air mixture. 



NACA 

RESEARCH ABSTRACTS NO. 93 



UNPUBLISHED PAPERS 



N-22024* 

FLOW NEAR A HEATED SOLID BODY IN A STAND- 
ING ACOUSTIC WAVE. P. N. Kubanskii. October 
1955. Up. diagrs. (Trans, of Zhurnal Tekhnich- 
eskoi Fiziki, v. 22, no. 4, April 1952, p. 585-592) 

Results of experiments carried out in standing waves 
generated by vibrations of finite amplitude are pre- 
sented. Two types of waves are considered; 
(1) forced standing waves, formed wtien the frequen- 
cy of outside vibrations does not correspond with the 
natural frequency of vibration of the radiating sys- 
tem, and (2) standing waves which form as the re- 
sult of coincidence of frequency of outside vibration 
with natural frequency of vibration of the radiating 
system. Optical methods are used in the examina- 
tion of the flow and the presence of higher harmonics 
in the standing acoustical wave is observed. The 
deformation of the waves is also studied. 



N-34452 

SOFT ROT. DESTRUCTION OF WOOD THROUGH 
COMMON FUNGI. (Moderfaule. Die Zersetzung 
von Holz durch niedere Pilze). W. P. K. Findlay and 
J. G. Savory. October 1955. 12p. photos., tab. 
(Trans, from Holz als Roh- und Werkstoff, v. 12, 
no.8, August 1954, p. 293-296). 

If wood is exposed to humid weather for some length 
of time, a softening of the surface takes place which 
is due to a fungus attack. This phenomenon is of 
minor significance except in the case of cooling 
towers and similar Industrial water-cooling plants. 
Soft rot, as this fungus is generally called, was pro- 
duced according to a method by Abrams. The fungi 
causing soft rot are reviewed and preservatives for 
protection of wood are discussed, but further inves- 
tigations are required in order to indicate effective 
methods of protection of wood from soft rot. 



N-39717'' 

EFFECT OF ACOUSTICAL VIBRATIONS OF FINITE 
AMPLITUDE ON THE BOUNDARY LAYER. P. N. 
Kubanskii. October 1955. lip. diagrs. (Trans, of 
Zhurnal Tekhnicheskoi Fiziki, v. 22, no. 4, April 1952, 
p.593-601) 

It is concluded that acoustical standing waves of 
finite amplitude can produce currents near the wall 
of a solid body, even when the body is located in a 
stream. A considerable intensity of acoustical 
vibration is necessary in order to produce acoustical 
flow near the walls of a solid body in a stream. The 
acoustical streams produced near the walls of solid 
bodies exert an influence on the boundary layer 
which surrounds the body. There is a possibility of 
using acoustical vibrations to control phenomena 
which occur in the boundary layer. Such phenomena 
could be directed toward the desired side of a physi- 
cal body, depending upon the condition of tlie bound- 
ary layer. 



DECLASSIFIED NACA REPORTS 



NACA RM A50J26a 

AERODYNAMIC CHARACTERISTICS INCLUDING 
PRESSURE DISTRIBUTIONS OF A FUSELAGE AND 
THREE COMBINATIONS OF THE FUSELAGE WITH 
SWEPT-BACK WINGS AT HIGH SUBSONIC SPEEDS. 
Fred B. Sutton and Andrew Martin. February 6, 
1951. 117p. diagrs., photos., tabs. 
(NACA RM A50J26a) 
(Declassified from Confidential, 10/14/55} 

As part of an NACA transonic research program, 
three sweptback wings with a fuselage were inves- 
tigated over a Mach number range from 0.40 to 0.94. 
These model wings had NACA 65A006 sections par- 
allel to the plane of symmetry. One of the model 
wings was swept back 35° and had an aspect ratio of 
6; the other two were swept back 45° and had aspect 
ratios of 4 and 6. Force and pitching-moment data, 
tabulated pressure measurements, downwash and 
dynamic-pressure characteristics, and tuft studies 
are presented. The approximate effects of wing 
elasticity on lift and moment data are also shown. 



THE FOLLOWING REPORTS HAVE BEEN 
DECLASSIFIED FROM CONFIDENTIAL 11/14/55, 
AND ARE UNAVAILABLE: 

RM L8K18a 
RM L52K24a 



THE FOLLOWING REPORTS HAVE BEEN 
DECLASSIFIED FROM CONFIDENTIAL, 11/14/55 



NACA RM A7K28 

HIGH-SPEED STABILITY AND CONTROL CHARAC- 
TERISTICS OF A FIGHTER AIRPLANE MODEL 
WITH A SWEPT-BACK WING AND TAIL. Charles 
P. Morrill, Jr., and Lee E. Boddy. April 14, 1948. 
47p. diagrs., photos. (NACA RM A7K28) 

Wind-tunnel tests were made at high subsonic Mach 
numbers of a model of a pursuit airplane with a 35° 
sweptback wing and tail. Data are included in the 
report which show the basic characteristics of the 
model; the control and hinge-moment characteristics 
of the horizontal tail, elevator, and aileron; the ef- 
fect of a wing leading-edge slat; and the effect of a 
fuselage-side dive brake. 



UNIVERSITY OF FLORIDA 



NACA 

RESEARCH ABSTRACTS NO. 93 



NACA RM A9K02 



3 1262 08153 293 8 

NACA RM A53C19 



INVESTIGATION OF DOWNWASH AND WAKE 
CHARACTERISTICS AT A MACH NUMBER OF 1.53. 
Ill - SWEPT WINGS. Edward W. Perkins and 
Thomas N. Canning. February 23, 1950. 41p. 
diagrs., tab. (NACA RM A9K02) 

The results of an experimental investigation of the 
downwash and wake characteristics behind two 
highly swept wings in a supersonic stream are pre- 
sented. The leading-edge sweep angles of the two 
wings were 63° and 63°45', the aspect ratios were 
3.50 and 1.66, and the corresponding taper ratios 
were 0.25 and 1.00. The tests were made at a Mach 
number of 1.53 and Reynolds numbers of 1.4 million 
and 2.6 million, respectively. A comparison be- 
tween experimental and theoretical values of the rate 
of change of downwash angle with angle of attack at 
zero lift is made. 



NACA RM A52K12 

TECHNIQUES FOR DETERMINING THRUST IN 
FLIGHT FOR AIRPLANES EQUIPPED WITH 
AFTERBURNERS. L. Stewart Rolls, C. Dewey 
Havill, and George R. Holden. January 1953. 27p. 
diagrs., photos. (NACA RM A52K12) 

An experimental technique has been developed which 
enables a determination of the net thrust for an 
afterburner -equipped airplane in flight. Measure- 
ments from a swinging pitot-static pressure and 
total temperature probe are used to determine the 
gross thrust, total air-flow rate, and net thrust. 
Details are also presented for an air-cooled fixed- 
pressure probe for the determination of basic engine 
thrust. 



NACA RM A52K13 

TESTS IN THE AMES 40- BY 80 -FOOT WIND TUN- 
NEL OF AN AIRPLANE MODEL WITH AN ASPECT 
RATIO 4 TRIANGULAR WING AND AN ALL- 
MOVABLE HORIZONTAL TAIL - HIGH-LIFT DE- 
VICES AND LATERAL CONTROLS. Ralph W. 
Franks. February 1953. 45p. diagrs., photo., 2 
tabs. (NACA RM A52K13) 

Tests have been made of a model consisting of a tri- 
angular wing in combination with a fuselage of fine- 
ness ratio 12.5; a thin, triangular, vertical tail with 
a constant -chord rudder; and a thin, unswept, all- 
movable horizontal tail. The wing had an NACA 
0005 modified section and was equipped with slotted 
inboard and plain outboard flaps. Tests were made 
with the wing -fuselage -vertical -tail configuration in 
addition to the tests of the complete model. The 
results of tests of lateral and directional controls, 
the inboard flaps as a high-lift device, and the out- 
board flaps as a high-lift device are presented. The 
Reynolds number, based on the wing mean aerody- 
namic chord, was appro.ximately 10.9 million and 
the Mach number was about 0.13. 



EXPERIMENTAL INVESTIGATION OF THE EF- 
FECTS OF PLAN-FORM TAPER ON THE AERO- 
DYNAMIC CHARACTERISTICS OF SYMMETRICAL 
UNSWEPT WINGS OF VARYING ASPECT RATIO. 
Edwin C. Allen. May 1953. 53p. diagrs., photos., 
tab. (NACA RM A53C19) 

This r.eport presents results of tests of a series of 
symmetrical, unswept, 8-percent-thick wings of 
varying aspect ratio and taper ratio.' The wings 
were tested in combination with four different bodies 
of revolution over a Mach number range from 0.40 
to 0.94 with a corresponding Reynolds number range 
from 2.58 million to 5.90 million. The lift, drag, 
and pitching-moment data are presented for wings 
of aspect ratios 2, 3, and 4 and for taper ratios of 
0.20 to 1.00. 



NACA RM E51K15 

COMPARISON OF LOCKED-ROTOR AND WIND- 
MILLING DRAG CHARACTERISTICS OF AN AXIAL- 
FLOW-COMPRESSOR TYPE TURBOJET ENGINE. 
K. R. Vincent, S. C. Huntley, and H. D. Wilsted. 
January 1952. lOp. diagrs. (NACA RM E51K15) 

The internal drag of an axial -flow turbojet engine 
with the rotor locked in place to prevent windmilling 
and with the engine windmilling was obtained over a 
range of simulated Mach numbers. The corrected 
internal drag of the engine with the locked rotor was 
210 pounds or only 46 percent of the windmilling 
drag at a flight Mach number of 0.8. 



NACA RM E53K06 

ANALYTICAL STUDY OF LOSSES AT OFF-DESIGN 
CONDITIONS FOR A FIXED-GEOMETRY TURBINE. 
Warner L. Stewart and David G. Evans. February 
1954. 48p. diagrs., tab. (NACA RM E53K06) 

An analytical investigation was made to determine 
the off-design loss characteristics of a fixed- 
geometry turbine of which the experimental per- 
formance was known. The method of analysis 
utilized an effective loss parameter and assumed 
that the velocity normal to the blade entrance angle 
was lost as a total -pressure loss. The method also 
assumed constant tangential component of velocity 
between the station just upstream and just down- 
stream of the stator and rotor trailing edge. Good 
correlation between the analytically and experi- 
mentally obtained performance was found over the 
entire map until limiting loading was approached. 
The large decrease in efficiency at low-speed high 
pressure ratios and at high-speed low pressure 
ratios was found in the analysis to be almost en- 
tirely due to the rotor incidence and turbine exit 
whirl losses. From the results of the investigation 
it was concluded that for turbines designed to oper- 
ate efficiently at more than one point, the design 
must compromise rotor incidence angle and exit 
whirl losses. 



NACA 
RESEARCH 



ABSTRACTS NO. 93 



NACA RM L8K12a 

HINGE -MOMENT MEASUREMENTS OF A WING 
WITH LEADING-EDGE AND TRAILING-EDGE 
FLAPS AT A MACH NUMBER OF 1. 93. William B. 
Boatright and Robert W. Rainey. January 14, 1949. 
12p. diagrs., tab. (NACA RM L8K12a) 

Hlnge-monnent data for a wing with leading-edge and 
tr'ailing-edge flaps of wedge section were obtained in 
the Langley 9 -inch supersonic tunnel at a Mach num- 
ber of 1.93 and a Reynolds number of 1.31 x 10^. 
Curves of hinge moment against angle of attack and 
against flap deflection are shown, and the results 
are compared with theory. The possibility of a 
linkage system to reduce control forces is discussed. 



NACA RM L8K17a 

CONTROL EFFECTIVENESS AND HINGE -MOMENT 
MEASUREMENTS AT A MACH NUMBER OF 1.9 OF 
A NOSE FLAP AND TRAILING-EDGE FLAP ON A 
HIGHLY TAPERED LOW-ASPECT-RATIO WING. 
D. William Conner and Meade H. Mitchell, Jr. 
January 10, 1949. 26p. diagrs., photo. 
(NACA RM L8K17a) 

Nose flaps and trailing-edge flaps were tested on a 
low-aspect-ratio, highly tapered, half-span wing 
model in the Langley 9- by 12-inch supersonic blow- 
down tunnel at a Mach number of 1.9 and a Reynolds 
number of 3,000,000. Lift, drag, pitching- and 
rolling-moment data for the wing and hinge-moment 
data for the flaps were obtained. All tests were 
made in the presence of a fuselage. 



NACA RM L8K24a 

EXPERIMENTAL AND CALCULATED HINGE 
MOMENTS OF TWO AILERONS ON A 42.7° SWEPT- 
BACK WING AT A MACH NUMBER OF 1.9. 
James C. Sivells and Kennith L. Goin. January 19, 
1949. 23p. diagrs., photos., tabs. 
(NACA RM L8K24a) 

A 42.7° sweptback wing was tested with two types of 
ailerons in the Langley 9- by 12-inch supersonic 
blowdown tunnel at a Mach number of 1.9 and a 
Reynolds number of 2.2 x 10^, The wing had an as- 
pect ratio of 4, a taper ratio of 0.5 , and an 8- 
percent-thick biconvex airfoil section. The contour 
of one aileron was formed by the basic airfoil con- 
tour, and the other aileron had flat sides and a 
trailing-edge thickness of one-half the hinge-line 
thickness. 




NACA RM L9K01a 

ROCKET-POWERED FLIGHT TEST OF A ROLL- 
STABILIZED SUPERSONIC MISSILE CONFIGURA- 
TION. Robert A. Gardiner and Jacob Zarovsky. 
January 12, 1950. 32p. diagrs., photos., tab. 
(NACARM L9K01a) 

A missile research model incorporafing wing-tip 
ailerons and a gyro-actuated automatic roll control 
was flight tested at supersonic speed. The aerody- 
namic rolling derivatives for zero-lift flight were 
determined from the flight record. It was concluded 
that the method used in the preflight system analysis 
is valid and that the gyro-actuateci4«r=attefQn control 
system provided a satisfactorv--l»«^a^J 
roll stabilization in zero-lift^s^ip^l^istJnicTTrghit'/^^^^ 



NACA RM L9K09 



AERODYNAMIC INVESTIGATI^JOF A PARABOLIC 
BODY OF REVOLUTION AT MAErff*mMBEI? OF 
1.92 AND SOME EFFECTS OF AN ANNULAR JET 
EXHAUSTING FROM THE BASE. Eugene S. Love. 
February 8, 1950. 75p. diagrs., photos., tab. 
(NACA RM L9K09) 

An aerodynamic investigation of a slender pointed 
parabolic body of revolution was conducted at a Mach 
number of 1.92 with and without the effects of an 
annular supersonic jet exhausting from the base. 
Measurements without the jet in operation were 
made of lift, drag, pitching moment, base pres- 
sures, and radial and axial pressures. With the jet 
in operation, pressure measurements were made 
over the rear of the body with the primary variables 
being angle of attack, ratio of jet velocity to stream 
velocity, and ratio of pressure at jet exit to stream 
pressure. 



NACA RM L50K06 

HORIZONTAL -TAIL EFFECTIVENESS AND DOWN- 
WASH SURVEYS FOR TWO 47.7° SWEPTBACK 
WING-FUSELAGE COMBINATIONS WITH ASPECT 
RATIOS OF 5.1 AND 6.0 AT A REYNOLDS NUMBER 
OF 6.0 X 10^. Reino J. Salmi. January 12, 1951. 
65p. diagrs., photos., 2 tabs. (NACA RM L50K06) 

Results of wind-tunnel tests on two 47.7° sweptback 
wing-fuselage combinations of aspect ratios 5.1 and 
6.0 to determine the effects of the vertical location 
of a horizontal sweptback tail on tail effectiveness 
and the static longitudinal stability were presented. 
The tests were made at a Reynolds number of about 
6.0 X 10^ (Mach number of 0.14) for various combi- 
nations of leading-edge and trailing-edge flaps. The 
results of airstream surveys in the region of the 
tail are also presented. 



10 



NACA RM L50K29 

LOW-SPEED LONGITUDINAL AND WAKE AIR- 
FLOW CHARACTERISTICS AT A REYNOLDS NUM- 
BER OF 6.0 X 106 OF A 52° SWEPTBACK WING 
EQUIPPED WITH VARIOUS SPANS OF LEADING- 
EDGE AND TRAILING-EDGE FLAPS, A FUSELAGE, 
AND A HORIZONTAL TAIL AT VAWOUS VERTICAL 
POSITIONS. Roland F. Griner and Gerald V. Foster. 
February 28, 1951. 66p. diagrs., photo., 3 tabs. 
(NACA RM L50K29) 

The results are presented of an investigation con- 
ducted in the Langley 19-ioot pressure tunnel at a 
Reynolds number of 6.0 x 10^ to determine the ef- 
fects of leading-edge-flap spans on an NACA 64- 
series wing swept back 52°. Several of the more 
satisfactory spans of leading-edge flaps were inves- 
tigated with various combinations of trailing-edge 
flaps, fences, a fuselage, and a horizontal tail. Sur- 
veys of downwash angle, sidewash angle, and dy- 
namic pressure ratio behind the wing at approxi- 
mately the location of a horizontal tail are presented. 



NACA RM L51I05 

EFFECT OF FORMATION POSITION ON LOAD 
FACTORS OBTAINED ON F2H AIRPLANES. Carl R. 
Huss and Harold A. Hamer. December 1951. 15p. 
diagrs., 3 tabs. (NACA RM L51I05) 

Plots of load factor against airplane position are 
presented for three combinations of four airplanes 
hying in formation. The plots show that the load- 
factor trend was to increase toward the end of the 
formation. Typical time histories are presented. 



NACA RML51K30 

TIME HISTORIES OF MANEUVERS PERFORMED 
WITH AN F-86A AIRPLANE DURING SQUADRON 
OPERATIONS. Harold A. Hamer and Campbell 
Henderson. February 1952. 90p. diagrs., 3 tabs. 
(NACA RM L51K30) 

Some preliminary results of maneuvers performed 
during U. S, Air Force squadron operations with an 
F-86A jet-fighter airplane are presented in time- 
history form. The maneuvers cover a speed range 
from the stall to 530-mile-per-hour indicated air- 
speed and pressure altitudes varying from sea level 
to approximately 25.000 feet. Variation of the 
maximum airplane linear and angular accelerations 
experienced during the investigation are also pre- 
sented. 



NACA 
RESEARCH 



ABSTRACTS NO. 93 



NACA RM L52K07 

FREE -FLIGHT INVESTIGATION AT ZERO UFT IN 
THE MACH NUMBER RANGE BETWEEN 0.7 AND 
1.4 TO DETERMINE THE EFFECTIVENESS OF AN 
INSET TAB AS A MEANS OF AERODYNAMICALLY 
RELIEVING AILERON HINGE MOMENTS. William 
M. Bland, Jr., and Edward T. Marley. January 
1953. 19p. diagrs., photos. (NACA RM L52K07) 

An experimental investigation employing a technique 
which utilized a zero-lift rocket -propelled model in 
free flight has been made to determine some of the 
characteristics of an inset tab as an aerodynamic 
balance in the Mach number range between 0.7 and 
1.4. The fixed, 0.09-chord, full-span, inset tab 
that was investigated was attached to a 0.3-chord 
full-span aileron on a wing of aspect ratio 3 and 
taper ratio 0.6 that had the quarter-chord line swept 
back 45° and NACA 65A006 airfoU sections parallel 
to the model center line. Results of this investiga- 
tion show that the tab was capable of balancing (trim- 
ming) the aileron hinge moments throughout the Mach 
number range investigated even though the effective- 
ness of the tab decreased with increasing Mach num- 
ber. It was shown that the aileron rolling effective- 
ness was decreased considerably when the tab was 
used to reduce the aileron hinge moments. The tab 
when considered as a servotab was an effective aero- 
dynamic balance for Mach numbers less than 1.1. 
At no time during the investigation did the mass- 
balanced aileron show any evidence of buzz or flutter. 
It was also shown that the tab effectiveness could be 
estimated with reasonable accuracy from e^qjerimen- 
tal data and from thin-airfoil theory. 



NACA RM L52K25 

INVESTIGATION OF THE EFFECT OF CHORDWISE 
POSITIONING AND SHAPE OF AN UNDERWING NA- 
CELLE ON THE HIGH-SPEED AERODYNAMIC 
CHARACTERISTICS OF A 45° SWEPTBACK 
TAPERED-IN-THICKNESS-RATIO WING OF AS- 
PECT RATIO 6. H. Norman Silvers and Thomas J. 
King, Jr. January 1953. 50p. diagrs. 
(NACA RM L52K25) 

An investigation at high ^eeds of chordwise posi- 
tioning of underwing nacelles at a spanwise location 
of 0.46 semispan with an ogive -cylinder shape 
(fineness ratio = 9.34), an NACA 65A-series airfoil 
of revolution shape (fineness ratio = 10.68), and a 
modified NACA 0-series airfoil of revolution shape 
reversed in direction (fineness ratio = 10.04) was 
made on a small-size 45° sweptback tapered-in- 
thickness-ratio wing of aspect ratio 6. 



NACA 
RESEARCH 



ABSTRACTS NO. 93 



NACA RM L53K16 

AN AIR-FLOW-DIRECTION PICKUP SUITABLE FOR 
TELEMETERING USE ON PILOTLESS AIRCRAFT. 
Wallace L. Ikard. March 1954. 25p. diagrs., 
photos. (NACA RM L53K16) 

A vane-type air-flow-direction pickup is described 
which is suitable for telemetering angle-of-attack 
and angle-of-sideslip data from rocket-propelled 
pilotless aircraft models. Test results which are 
presented show that the device performs well under 
high accelerations and is stable throughout a Mach 
number range from subsonic to above a Mach num- 
ber of 2.5. 



NACA RM L53K18 

EXPERIMENTAL INVESTIGATION OF THE OSCIL- 
LATING FORCES AND MOMENTS ON A TWO- 
DIMENSIONAL WING EQUIPPED WITH AN OSCIL- 
LATING CIRCULAR-ARC SPOILER. Sherman A. 
Clevenson and John E. Tomassoni. January 1954. 
20p. diagrs., photos. {NACA RM L53K18) 

Results of a wind-tunnel investigation of the forces, 
moments, and phase angles on a two-dimensional 
wing equipped with an oscillating circular -arc 
spoiler are presented. Schlieren photographs are 
presented which show the flow over and behind the 
spoiler. Data for Reynolds numbers from 1.3 x 10° 
to 6.3 X 10^, Mach numbers from 0.2 to 0.82, and 
reduced frequencies from to 0.92 on the normal- 
force and pitching-moment coefficients and their 
respective phase angles referred to spoiler position 
are indicated. 



NACA RM L53K30 

PRELIMINARY INVESTIGATION OF THE FLOW IN 
AN ANNULAR-DIFFUSER— TAILPIPE COMBINA- 
TION WITH AN ABRUPT AREA EXPANSION AND 
SUCTION, INJECTION, AND VORTEX-GENERATOR 
FLOW CONTROLS. John R. Henry and Stafford W. 
Wilbur. February 1954. 27p. diagrs. 
(NACA RM L53K30) 

The performance of an annular-diffuser — tailpipe 
combination with an abrupt area expansion was in- 
vestigated with and without flow controls in the form 
of suction, injection, and vortex generators. The 
diffuser had a 21 -inch-diameter straight outer wall, 
an area ratio of 1.9 to 1, and fully developed pipe 
flow at the inlet. Inlet Mach number was varied be- 
tween 0.18 and 0.43. The ratio of the auxiliary air 
flow to the flow of the main stream was varied from 
to approximately 4 percent. (Both suction and 
injection flow controls were effective in producing 
improved diffuser performance. ) 



NACA - Langley Field. Vj.